1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine airfoils with cooling circuits.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine produces mechanical power from burning a fuel. A compressor supplies compressed air to a combustor in which a fuel is burned to produce an extremely hot gas flow. The hot gas flow is passed through a turbine to convert the hot gas flow into mechanical energy by driving the turbine. In a typical industrial gas turbine, the turbine shaft drives the compressor and an electric generator to produce electrical power.
The engine efficiency can be increased by providing for a higher temperature in the hot gas flow entering the turbine. An industrial gas turbine typically has four stages with stator vanes located upstream of the rotor blades. The first stage stator vanes and rotor blades are exposed to the highest flow temperature. Therefore, the materials used in this turbine parts limit how high the temperature can be.
Once the materials used for the first stage vanes and blades are maximized with respect to the highest temperature allowed, the airfoils in question can include cooling air to allow for a further increase in the operating temperature. Complex cooling air circuitry has been proposed in the prior art to not only maximize the cooling ability of the airfoils but to also minimize the amount of cooling air sued. Since the pressurized cooling air used in these airfoils typical comes from bleed off air from the compressor, minimizing the amount of cooling air used will also increase the efficiency of the engine.
Turbine airfoils generally include hot spots on the airfoil surface where higher temperatures are found. Therefore, some parts of the airfoil require more cooling than other parts. Hot spots can reduce the life of a turbine airfoil due to lack of adequate cooling in the certain spots. Turbine airfoils can be cooled by a combination of convection cooling, impingement cooling, and film cooling. One or more for these cooling methods can be used in selective locations around the airfoil.
Another way in which the increased use of cooling air can be avoided, or cooling air requirements can be reduced, is by providing metal parts that are capable of operating above the maximum use temperature of 1,150.degree. C. The provision of metal parts capable of operating at temperatures beyond 1,150.degree. C. would allow either relaxation of cooling requirement or the reduction or elimination of the dependence on the thermal barrier coatings, or both.
It is also well known that the operating efficiency of gas turbine engines may be improved by reducing the total weight of the metal parts utilized. Currently, because of the required intricate internal cooling passages within metal parts such as blades and vanes, particularly near their outer surfaces, and the fragile nature of the ceramic cores used to define these passages during formation, it is necessary to utilize large tolerances that allow for the possibility of core shifting. The use of materials and processes that would simplify the design requirements for these internal passages would permit the amount of material used in each metal part to be reduced. Also, the use of materials that are less dense would achieve weight reductions for each metal part. Small savings can be significant because of the large number of these metal parts that are utilized in a typical engine.
Prior art near wall cooling arrangements utilized in an airfoil main body is constructed with radial flow channel plus re-supply holes in conjunction with film discharge cooling holes. As a result of this cooling construction approach, span-wise and chord-wise cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. In addition, a single radial channel flow is not the best method of utilizing cooling air. This results in a low convective cooling effectiveness.
U.S. Pat. No. 5,640,767 issued to Jackson et al on Jun. 24, 1997 and entitled METHOD FOR MAKING A DOUBLE-WALL AIRFOIL discloses a turbine airfoil with an airfoil skin formed over a partially hollow airfoil support wall with a plurality of longitudinally extending internal channels formed between the skin and the wall. Film cooling holes are formed in the skin after the skin has been secured to the airfoil wall. Because the internal channels that supply cooling air extend along the span-wise length of the airfoil, the cooling requirements cannot be adjusted for along the span-wise direction of the airfoil.
U.S. Pat. No. 6,582,194 B1 issued to Birkner et al on Jun. 24, 2003 and entitled GAS-TURBINE BLADE AND METHOD OF MANUFACTURING A GAS-TURBINE BLADE discloses a turbine blade with a metal blade body having peg-like elevations extending outward and forming spaces between adjacent pegs. A coating of ceramic material is applied within the spaces and flush with a top of the pegs, and a covering coat applied over to form an outer wall of the blade. The ceramic material is leached away to leave impingement spaces. Oblique film cooling holes are then formed in the outer wall. The impingement channels in the Birkner patent also extends along the span-wise length of the blade, and therefore the amount of cooling air cannot be adjusted to vary the cooling amount along the span-wise direction of the blade. In the above cited prior art references, the film cooling holes are formed in the outer wall of the airfoil in a separate process, usually by laser drilling the holes. Drilling the film cooling holes after the outer wall and the impingement cell or cavity has been formed requires an extra manufacturing process that increases the cost of making the airfoil.
Thin walled airfoils are desirable because the thin walls can be cooled by impingement air and film cooling air. However, thin walls are difficult if not impossible to cast into an airfoil. An improvement for the airfoil main body near-wall cooling can be achieved by incorporation of the present invention into the airfoil main body cooling design of the cited prior art references.
It is an object of the present invention to provide for a turbine airfoil with a thin outer wall having near wall cooling.
It is another object of the present invention to provide for a turbine airfoil with cast in place film cooling holes in order to reduce the manufacture steps to make the airfoil.
It is another object of the present invention to provide for a turbine airfoil that includes a plurality of modules to provide cooling to the airfoil at pre-specified amounts in order to increase the life of the airfoil.